Rotorcraft with interchangeable rotor diameters

ABSTRACT

A tiltrotor aircraft is designed to accommodate rotors of different diameters, as well as corresponding wings and fuselages with different span and length, while maintaining very high parts commonality, especially with respect to drive train and power source. This enables design and operation of a fleet of such aircraft with significantly different rotor diameters, which are nevertheless optimized for different missions. In preferred embodiments the rotors are configured to have high stiffness and low weight to reduce aero-structural dynamic issues across the fleet. Also in preferred embodiments drive systems are designed for a full range of speed, torque, and power associated with all intended rotors. Turboshaft engine speeds are restricted to a narrow RPM range, so that a single gearset can be replaced to achieve the desired rotor RPM. Also in preferred embodiments, aircraft in a fleet can differ in folded length, empty weight, payload length by up 50%.

FIELD OF THE INVENTION

The field of the invention is a tiltrotor aircraft.

BACKGROUND

The following description includes information that may be useful inunderstanding the currently disclosed subject matter. It is not anadmission that any of the information provided herein is prior art orrelevant to the presently claimed invention, or that any publicationspecifically or implicitly referenced is prior art.

Tiltrotor aircraft are a unique type of aircraft which are capable ofchanging geometry in flight to provide hover capability and wingbornecruise flight in the same platform. Prior art consists primarily oftiltrotor aircraft designs which have existed only as limited productionexperimental demonstration vehicles. Two exceptions exist: the BellBoeing V-22 for military applications with a fleet size of 375 aircraftand the AgustaWestland AW-609 which is currently undergoing civilian andmilitary certification. While these tiltrotors, like most aircraft, aresuited to a variety of missions, the rotor diameter of each design iskept constant for all missions and variants.

Most aircraft development programs capitalize on multiple customers bycreating variants of a baseline design best suited to each customer'srequirements. Prior art covers many instances of components such asmission equipment and engines installed on the aircraft variants. Anexample is the two variants of the Sikorsky S-70: A ship-based MR-60RSeahawk® equipped for anti-submarine/anti-surface warfare, and thecivilian S-70 Firehawk® for firefighting. Fuselage variants also exist,typically by adding a “plug” to lengthen the fuselage as on the AirbusA320 variants. Seehttps://upload.wikimedia.org/wikipedia/commons/9/92/A32XFAMILYv1.0.png.Significant commonality can be maintained, and consequently developmentand recuring costs are minimized by anticipating these variants in theinitial design.

A specific example of the desire for aircraft variants is the shareddesire of the Army and Navy to develop a new vertical takeoff andlanding (VTOL) aircraft. While many requirements, such as carryingtroops and weapons at high speed are shared, some key differenceschallenge the goal of commonality. A large rotor diameter is desired forbest hover capability. The Army flies from open fields and runways whichdo not constrain the rotor diameter, so a large rotor diameter isoptimal. The Navy requires its aircraft to takeoff, land, and stowonboard its ships. Clearances to superstructure and other aircraft limitthe allowable rotor diameter. It would therefore be beneficial toprovide variations of a tiltrotor aircraft with different rotordiameters.

However, a change in the diameter of a lifting rotor, such as on atiltrotor aircraft, is not accommodated for in baseline designs due tothe challenge of dynamics and vibration which require precise design andthorough evaluation. Additionally, a conventional tiltrotor aircraftwith a fully articulated or semi-rigid rotor is limited in maximumflight speed by the onset of whirl flutter. This is characterized byunstable aero-structural interaction between the rotor and the wing.Modifications to the wing or rotor, especially wingspan and rotordiameter, of a conventional tiltrotor aircraft significantly impactwhirl flutter characteristics. A change in the required rotor diameterinvokes a change in inboard wingspan to preserve fuselage to rotorclearance. This motivates a full redesign and reevaluation of theaircraft including its propulsion system.

A person of ordinary skill in the art attempting to satisfy the Army andNavy requirements would choose a single rotor diameter allowed by bothservices and therefore compromise the potential benefits of a differentrotor design for each service. This happened in the JVX program, theArmy-Navy joint program which eventually produced the V-22, the onlymass-produced tiltrotor in existence. The V-22's rotor is about fivefeet smaller in diameter than ideal for the Army requirements asdescribed in the following quote:

“V-22 rotor diameter was constrained by shipboard operation. Required totaxi past the superstructure of such ships with its rotors no less than12 feet 8 inches (3.9 m) away from the “island” and its outboard tiresat least five feet inboard from the edge of the deck, the V-22'sproprotors could be no more than 38 ft (12 m) in diameter—about fivefeet less than ideal for an aircraft that size, according to engineersworking on the project at the time.”https://vtol.org/files/dmfile/JMR_Bell-Vertiflite.pdf

One approach to improve efficiency of a tiltrotor is an in-flightvarying of rotor diameter. The narrow range of engine RPM motivates avariable rotor diameter. A larger diameter setting, and therefore largerdisk area, is used in hover when high rotor thrust is required. Diameteris reduced for wingborne cruise when thrust demand is lower. Multipleinstances of prior art describe a variable diameter rotor which allowedvarying rotor disk area in flight. US Patent USD401898S describes theaircraft envisioned by Sikorsky Aircraft Corp using a Variable DiameterTiltrotor (VDTR) in the 1990s. This prior art was aimed exclusively atproviding a large rotor diameter for hover and for rotor edgewise flight(like a helicopter) and for reduced rotor diameter to allow tilting andfuselage clearance. While Sikorsky did ground testing of VDTR subsystemsthe high rotor system complexity and flight safety risks madeapplications of VDTR undesirable.

Additional research by Farhan Gandhi at Penn State University describesa variable diameter tiltrotor aircraft where centrifugal forces withinthe rotor control the diameter, meaning rotor diameter and rotationalspeed (RPM) are proportional. See Gandhi, Farhan, “Length-Morphing RotorReady to Provide Helicopter Versatility”, Popular Mechanics, Oct. 1,2007 Breakthrough Awards. This aircraft is designed for the full rangeof rotor diameters of the rotor system.

A particular challenge of this concept is the extremely large forcerequired to control the blade extension and retraction. The diametercontrol mechanism and duplicate structure associated with an extendablerotor blade prohibit a lightweight, stiff rotor as contemplated herein.Consequently, these examples of prior art did not include aninterchangeable rotor with different diameter. U.S. Pat. No. 9,045,226covers aircraft variants which change fuselage modules, wingconfigurations, and rotor configurations. The rotors are extendable in atelescoping manner like those in the Sikorsky patent. This feature isused to reduce the ground footprint of the vehicle when the rotor isstopped.

Prior art covers an extension of the structure connecting two rotors.Erciyes University describes a multicopter drone system where armsextending to the rotors are exchanged for different lengths. Whilechanges such as arm length or rotor diameter on a small scale (approx.20 kg and 18 inch rotor diameter) multicopter are commonplace, thisbecomes challenging at larger scales described in the present patent.Rotor development and evaluation for rotor diameters in the range of 12to 80 feet or more, described herein, take years of analysis and testingfor each configuration.

Prior art includes helicopters with rotor systems which have a differentblade count. The MD Helicopters MD520N and MD600 are an example of this.The MD520N has a 5 bladed main rotor. The maximum takeoff weight, power,and fuselage size were increased for the MD600. This motivated a higherblade count rotor with 6 blades. An additional blade does notsignificantly impact the dynamic characteristics of the blade, and theapplication to a helicopter rather than a tiltrotor aircraft meanscoupled dynamics with a wing structure are not considered. These rotorsystems are not interchangeable between models, and the other aircraftsystems such as the engine and transmission are specific to each rotortype.

Other examples of interchangeable rotor geometries exist. U.S. Pat. No.7,246,998 B2 describes a helicopter rotor with a replaceable tipsegment. It suggests a replaceable tip segment from 87% radius to thetip. The inventive subject matter replaces the entirety of the rotorblade, such that the characteristic dimensions, including rotordiameters, are significantly different after replacement.

Some developmental rotorcraft featured differing rotor diameters whichwere introduced to solve dynamic instabilities or modify performance. Anexample of a tiltrotor is the Bell XV-3 of 1955-1962. Initial rotors ofXV-3 had three fully articulated blades per rotor with 25-foot diameter.Serious flutter problems with the three bladed rotors in the wind tunnelmotivated replacing them with two bladed, semi-rigid rotors with 24-footdiameter. See http://www.aviastar.org/helicopters_eng/bell_xv-3.php.

While propeller aircraft may be certified for multiple interchangeablepropeller options, each propeller type must be certified to the FAA witha Supplemental Type Certificate. “An STC will probably be required if asignificant amount of analysis or flight tests are required, or ifextensive flight manual changes are necessary.” See FAA AC 21-40 Eventhough it is known for fixed-wing aircraft to change propellertypes/sizes, doing so with a rotorcraft is much more challenging becausenatural frequencies of large rotors and wing structures areconventionally closely spaced, resulting in whirl flutter concerns.

The prior art teaches lightweight stiff rotor blades. The low mass andhigh stiffness of such blades mitigate the aero-structural concernsassociated with traditional rotors. The behavior of high stiffnessrotors is predictable and critical frequencies of the wing and rotor areseparated to avoid interaction. A stiff rotor system can vary itsrotational speed (RPM) in flight to provide best efficiency withoutcompromising hover performance. This enables large rotor diameter rotorsto be used for efficient wingborne flight on a tiltrotor. See U.S. Pat.No. 6,007,298 (Karem, OSR), and U.S. Pat. No. 6,641,365 (Karem, OSTR).

The '298 and '365 patents, as well as all other extrinsic materialsdiscussed herein are incorporated by reference to the same extent as ifeach such materials were specifically and individually indicated to beincorporated by reference. Where a definition or use of a term in anincorporated reference is inconsistent or contrary to the definition ofthat term provided herein, the definition of that term provided hereinapplies and the definition of that term in the reference does not apply.

SUMMARY OF THE INVENTION

The inventive subject matter provides apparatus, systems and methods inwhich an aircraft with tilting rotors is designed to accept rotors ofdifferent diameters, as well as corresponding wings and fuselages withdifferent span and length, respectively.

Aircraft variants designed for specific missions can each benefit from aspecific rotor diameter. A small rotor can be motivated by maximumallowable folded dimensions for storage onboard a ship. A large rotor ismotivated by carrying heavier payloads or more fuel with the same power.A preferred embodiment has replaceable rotors that provide significantlydifferent rotor diameters for different missions, while maintainingcommonality in most aircraft systems, including the powerplant,drivetrain, and structural interfaces.

Drive systems contemplated herein are designed for the full range ofspeed, torque, and power associated with all intended rotors. Turboshaftengine speeds are restricted to a narrow range, so a single gearset isreplaced to achieve the desired rotor RPM. Except for the singlereplaced gearset, the drivetrain in each nacelle is common betweenvariants.

The inboard wingspan between nacelles is extended for a larger rotor tomaintain acceptable rotor to fuselage clearance. Interfaces from wing tonacelle and from wing to fuselage are common between variants. Thisallows interchangeable wing and nacelle systems with different diameterrotors.

If payload density is consistent between variants, a larger fuselagepayload volume is desired for a variant with higher payload weightcapacity. A fuselage extension of the constant cross-section segmentprovides increased fuselage volume. The wing to fuselage interface iscommon between variants, and fuselage internal features such as tiedowns or seats are replicated in the extended portion of the fuselage.

BRIEF DESCRIPTION OF THE DRAWING

FIG. 1 is a perspective view of a preferred aircraft in wingborne flightconfiguration.

FIGS. 2A-2C are perspective views of the aircraft of FIG. 1, showing arange of nacelle tilt orientations.

FIG. 3 is a top view section drawing of the nacelle and drivetrain ofFIG. 1.

FIG. 4A is a front view of the aircraft of FIG. 1, in wingborne flightconfiguration.

FIG. 4B is a top view drawing of the aircraft of FIG. 1, in wingborneflight configuration.

FIG. 5A is a perspective view of the aircraft of FIG. 1, in its foldedconfiguration.

FIG. 5B is a perspective view of a variant aircraft having a largerrotor diameter, longer wingspan, and longer fuselage, in its foldedconfiguration.

FIG. 6A is a front view of the aircraft of FIG. 1, in its unfoldedconfiguration.

FIG. 6B is a front view of the variant aircraft of FIG. 5B, in itsunfolded configuration.

FIG. 7A-7B are side views of the aircrafts of FIGS. 6A and 6B,respectively.

DETAILED DESCRIPTION

The following discussion provides many example embodiments of theinventive subject matter. Although each embodiment represents a singlecombination of inventive elements, the inventive subject matter isconsidered to include all possible combinations of the disclosedelements. Thus if one embodiment comprises elements A, B, and C, and asecond embodiment comprises elements B and D, then the inventive subjectmatter is also considered to include other remaining combinations of A,B, C, or D, even if not explicitly disclosed.

As used herein, and unless the context dictates otherwise, the term“coupled to” is intended to include both direct coupling (in which twoelements that are coupled to each other contact each other) and indirectcoupling (in which at least one additional element is located betweenthe two elements). Therefore, the terms “coupled to” and “coupled with”are used synonymously.

The inventive subject matter provides apparatus, systems, and methods inwhich a tiltrotor aircraft is designed to accommodate interchangeablerotors of dissimilar diameters. Preferred embodiments include a fleet ofat least two of such aircraft, each with rotors with differentdiameters. More preferred embodiments of fleets contain three, four,five or more of such aircraft.

FIG. 1 depicts a preferred aircraft, 100, in the wingborne mode ofoperation, comprising a wing 130, a fuselage 140, and first and secondrotors 110 comprising multiple blades 111. The aircraft 100 issubstantially symmetrical about the longitudinal centerline, such that,other than possibly being mirror images, the first and second rotors 110and their respective blades are substantially identical.

The rotor blades are of a stiff hingeless type such as that described inU.S. Pat. No. 6,641,365. First rotor 110 rotates about axis 112 togenerate thrust generally along the axis 112 and moments. First rotor110 is mounted to a tilting nacelle 120 which rotates around a tilt axis122 allowing operation across the range of rotorborne mode with rotorthrust pointed vertically through wingborne mode with the rotor thrustpointed forward. An outboard wing 131 is attached to a tilting nacelle120. Outboard wing control surface 134 provides roll control and reduceswing loads in transition/conversion flight mode. An inboard wing 132connects nacelle 120 to fuselage 140. It includes control surfaces 133for controlling the attitude of the vehicle and reducing the download inhover. The fuselage 140 envisioned for a manned configuration includes aforward cockpit 141 and cabin doors 142. Sponsons 143 on each side ofthe fuselage 140 feature doors 144 which allow access to the sponsonvolume and provide surfaces for mounting internally carried payloads.Tail surfaces 145 are attached to the fuselage 140 and provide vehicleattitude control primarily in wingborne flight. The aircraft 100 issubstantially symmetrical about the centerline, such that other thanpossibly being mirror images, nacelles 120, wings 130, and fuselage 140features, are substantially identical.

FIGS. 2A-2C depict the range of nacelle tilt angles. FIG. 2A shows theaircraft 100 in wingborne mode operation. FIG. 2B shows the aircraft 100in the intermediate transition/conversion mode operation. FIG. 2C showsthe aircraft 100 in rotorborne mode operation.

FIG. 3 shows a top view cross section view of nacelle 120 in wingborneorientation. The nacelle 120 contains the propulsion system comprisingan engine 321, and various drivetrain elements: Output speed determininggearset 322, tilt-axis gearbox 323, cross wing driveshaft 324 and thereduction gearbox 325. The cross wing driveshaft 324 is aligned insidethe inboard wing 132. The inboard wing 132 includes an inboard wing flap133. The outboard wing 131 is attached to the nacelle and includes anoutboard wing control surface 134.

Rotors of different diameter have different design rotational speeds,however turboshaft engines are limited to a narrow range of RPM. Theoutput speed determining gearset 322 is a single gearset in the nacelledrivetrain which can be replaced for variant aircraft to maintain thedesired gear ratio of engine RPM to rotor RPM. The tilt axis gearbox andcross wing driveshaft allow power transfer from one tilting nacelleacross the wing to the other nacelle and rotor system. This is criticalfor flight safety in an engine out condition.

The reduction gearbox 325 can contain a speed changing mechanism whichallows selection of different rotor speeds for different flightconditions. The reduction gearbox 325 transmits torque to the rotorblades through a hub structure 326. Common interfaces 327 of rotorblades 111 to the hub structure 326 enable interchangeability of rotorswith different diameter on variant aircraft

FIG. 4A depicts a front view of aircraft 100 in wingborne modeoperation. Rotor separation dimension 412 also defines the inboardwingspan. The circle swept by the blade tip of rotor 110 defines therotor diameter dimension 413. First aircraft 100 has a rotor separationdimension 412 of 38 feet and a rotor diameter 413 of 29 feet.

FIG. 4B depicts a top view of aircraft 100 in wingborne mode operationand depicts the fuselage length dimension 411. The first aircraft 100has a fuselage length dimension 411 of 41.8 feet. Rotor separationdimension 412 is also shown in this view.

An especially important concern for a naval operator is the dimensionsof the aircraft when on a ship. FIG. 5A depicts aircraft 100 folded tominimize its dimensions onboard a ship. Limited deck space constrainsthe width and length of an aircraft on a ship. The height of theaircraft may also be limited by an internal hangar dimension. Aircraft100 folds various components to minimize its dimensions: outboard wings131 with control surfaces 134 fold down to reduce total wingspan, rotorblades 111 fold toward the centerline, the entire wing/nacelle/rotorsystem rotates above the fuselage 140 to align the wingspan with thefuselage longitudinal direction, and tail surfaces 145 fold alongsidethe fuselage 140. Landing gear 146 are independently adjustable inheight to minimize folded height while also providing necessary groundclearance during landing and loading. Aircraft 100 has a folded lengthdimension 551 of 42.5 feet.

FIGS. 5B, 6B, and 7B show a second aircraft 500, not drawn to scale, butwhich should be interpreted with dimensions described herein. Secondaircraft 500 has rotors 510 with rotor diameter dimension 562 of 36feet. Its rotor separation dimension 561 is 45 feet. Aircraft 500 has afolded length dimension 552 of 49.5 feet.

FIG. 5B shows second aircraft 500 in its folded configuration. Secondaircraft 500 comprises a wing 530, a fuselage, 540, and a first andsecond rotor 510 comprising multiple blades 511. Rotors are attached tonacelles 520. Outboard wings 531 with control surfaces 534 are alsoattached to nacelles 520. Inboard wings 532 with control surfaces 533connect nacelles 520 to fuselage 540. The fuselage 540 envisioned for amanned configuration includes a forward cockpit 541 and cabin doors 542.Sponsons 543 on each side of the fuselage 540 feature doors 544 whichallow access to the sponson volume and provide surfaces for mountinginternally carried payloads. Tail surfaces 545 are attached to thefuselage 540 and provide vehicle attitude control primarily in wingborneflight. The aircraft 500 is substantially symmetrical about thecenterline in wingborne configuration, such that other than possiblybeing mirror images, rotors 510, nacelles 520, wings 530, and fuselage540 features, are substantially identical.

FIG. 6A shows the first aircraft 100 having a rotor 110 with rotordiameter 412 of 29 feet.

FIG. 7A shows the first aircraft 100 with fuselage length dimension 410of 41.8 feet. The internal payload bay extending from behind the cockpit141 to line 711 has a length dimension 701 of 17 feet. FIG. 7B shows thesecond aircraft 500 with fuselage length dimension 710 of 48.8 feet. Theinternal payload bay of aircraft 500 extending from behind the cockpit541 to line 712 has length dimension 702 of 24 feet.

Aircraft 100 and 500 have at least 80% commonality by part count of the“green aircraft”. Green aircraft in the aero industry vernacular is thecomplete flying aircraft excluding parts of the aircraft not requiredfor flight such as cabin furnishing, external paint and mission systems.In the case of a military aircraft the mission systems include theweapons system, sensor equipment and other mission-specific equipment.The level of part commonality is therefore maintained in the airframe,airframe folding (if so equipped), icing and lightning protection,propulsion, flight controls, landing gear, fuel system, electrical powersystem, lighting, cabin pressurization and air-conditioning, basiccockpit sensors and instruments, cockpit and cabin doors, windows andglazing. Interfaces between major components such as the wing tofuselage, wing to nacelle, and rotor blades to nacelles are commonbetween variants.

As used herein, “percentage commonality by part count” refers only toparts that weigh more than one pound. This limitation is intended toprevent a potential competitor from circumventing the claims by adding alarge number of tiny components such as ball bearings.

Viewed from another perspective, aircraft 100 and 500 have at least 80%commonality in power source and drivetrain, and more preferably at least85% commonality in power source and drivetrain, and even more preferablyat least 90% commonality in power source and drivetrain. In mostpreferred embodiments, the un-commonality in the power source anddrivetrain is substantially limited to a single gearset in each nacelle.

Rotor rotational speed (RPM) of most rotorcraft is preferred to be highto carry the most lift in hover but is constrained by a maximum tipspeed remaining below the speed of sound in all modes of flight. Therotor tip speed is a product of rotor RPM and rotor diameter. Tomaintain similar tip speeds required for lift in hover an aircraft witha smaller rotor requires a proportionally higher RPM. However, theoutput speed (RPM) of a conventional turboshaft engine at maximum poweris constant. In a preferred embodiment of the aircraft, a single gearset322 is replaced to maintain the desired gear ratio of engine RPM torotor RPM for rotors of different diameters.

The aircraft according to teachings herein have rotor diameters in therange of 12 to 80 feet, inclusive. This range limitation excludes smallaircraft which do not face the same rotor aero-structural dynamicschallenges associated with large aircraft. Even larger rotors arecontemplated, up to 90 or even 100 feet, and the drawing figures shouldbe interpreted accordingly.

A minimum 10% difference of rotor diameter represents a significantchange in rotor disk area and is meant to exclude minor modificationssuch as an exchangeable tip. Conventional rotor systems would require areevaluation of the rotor system and drivetrain to accommodate rotordiameter modifications in excess of 10%. Even more significant changesin rotor diameters of 15%, 20% or even 24% are contemplated. As usedherein, differences are measured from the lower number. For example, thedifference between the 29 foot rotor diameter of FIG. 5A and the 36 footrotor diameter of FIG. 5B is 24%. In contrast, conventional tiltrotordesigns would require a redesign of the rotor system, drive system andstructure to provide for aero-structural stability and to acceptincreased loads and different whirl flutter characteristics.

Rotors contemplated herein can be constructed using techniquesidentified in the above-referenced patents, U.S. Pat. No. 6,007,298(Karem, OSR), and U.S. Pat. No. 6,641,365 (Karem, OSTR). Keycharacteristics are high stiffness and light weight. In accordance withthe teachings of these two patents, aircraft contemplated herein wouldhave total weight of each blade in lbs. that does not exceed 0.0015times the diameter of the rotor in feet cubed. Flap stiffness of theblades, measured at 10% of the rotor radius, in lbs-in2 is not less than25 times the diameter of the rotor in feet to the fourth power.

The combination of high stiffness and light weight characteristicsavoids the problematic interaction of the blade's natural flap, lag, andtorsional oscillation with the rotor excitation frequencies as describedin the '365 patent. An exemplary embodiment described in the '365 patentutilizes carbon-epoxy advanced composite material to provide the highstiffness to weight ratio required. The blades are attached to a hub ina hingeless configuration, meaning there is no flap or lead-lagarticulation at the blade root. The hingeless configuration maintainsthe required stiffness of the rotor system and allows the transfer oflarge moments from the rotor system to the aircraft structure formaneuvering.

When designing a fleet of aircraft with substantially different rotordiameters, but otherwise a very high level of commonality, additionalbenefits result from the use of lightweight rigid rotors. Continuousoperation across a wide envelope of RPM, including for example 40% to100% of maximum rotor RPM, allows optimal rotor RPM for flight at allairspeeds. This is contemplated to allow sustained in-flight operationwith at least 25% reduction in RPM between hover and forward speed.

FIG. 14-20 of the '365 patent illustrate the beneficial power reductionand improved propeller efficiency associated with optimum speedtiltrotors compared to conventional tiltrotors. These benefits areoperationally relevant to commercial and military applications becausethey reduce vehicle weight, cost, and fuel consumption for a givenmission set. A multi-speed transmission can be used to provide fullrotor RPM range while limited to the narrow RPM range of a turboshaftengine.

It is contemplated that a fleet of aircraft according to teachingsherein include at least two aircraft, wherein a smallest and largestfolded length dimensions of the aircraft differ by between 15% and 50%,inclusive. The folded length of the aircraft can be a result of thewingspan, rotor diameter, fuselage length, or other components whichprotrude from the aircraft's basic geometry. The range of difference infolded length is intended to exclude minor modifications which affectthe folded length. The range is also limited to a difference in overallsize where a common drivetrain would be inefficient or ineffective.

In a contemplated fleet, differences between smallest and largestmanufacturers empty weight of aircraft can differ by between 15% and 50%of the first aircraft, inclusive. The range of difference in emptyweight is intended to exclude minor modifications such as constructionmethod or additional features. The range is also limited to a differencein overall size where a common drivetrain would be inefficient orineffective.

Also in a contemplated fleet, differences between smallest and largestpayload weight capacity can differ by between 500 and 50,000 pounds,inclusive. This range of difference in payload weight is also intendedto exclude minor modifications of the rotor, propulsion, and fuselagewhich affect payload weight capacity. Rotor lift capability to powerratio is generally related to rotor diameter squared. Therefore, thepreferred rotor diameter differences result in large lift and payloadweight differences. A difference in fuselage length of the preferredembodiments accommodates the difference in payload weight capacity aswell as differences in payload density relevant to different users. Forexample, carrying lower density payload such as soldiers requires alonger fuselage length compared to the same weight of a higher densitypayload such as water. In a contemplated fleet, the largest payloadlength dimension of the first and second aircraft differ by between 15%and 50%, inclusive.

In some embodiments, the numbers expressing quantities of ingredients,properties such as concentration, reaction conditions, and so forth,used to describe and claim certain embodiments of the invention are tobe understood as being modified in some instances by the term “about.”Accordingly, in some embodiments, the numerical parameters set forth inthe written description and attached claims are approximations that canvary depending upon the desired properties sought to be obtained by aparticular embodiment. In some embodiments, the numerical parametersshould be construed in light of the number of reported significantdigits and by applying ordinary rounding techniques. Notwithstandingthat the numerical ranges and parameters setting forth the broad scopeof some embodiments of the invention are approximations, the numericalvalues set forth in the specific examples are reported as precisely aspracticable. The numerical values presented in some embodiments of theinvention may contain certain errors necessarily resulting from thestandard deviation found in their respective testing measurements.

As used in the description herein and throughout the claims that follow,the meaning of “a,” “an,” and “the” includes plural reference unless thecontext clearly dictates otherwise. Also, as used in the descriptionherein, the meaning of “in” includes “in” and “on” unless the contextclearly dictates otherwise.

The recitation of ranges of values herein is merely intended to serve asa shorthand method of referring individually to each separate valuefalling within the range. Unless otherwise indicated herein, eachindividual value is incorporated into the specification as if it wereindividually recited herein. All methods described herein can beperformed in any suitable order unless otherwise indicated herein orotherwise clearly contradicted by context. The use of any and allexamples, or exemplary language (e.g. “such as”) provided with respectto certain embodiments herein is intended merely to better illuminatethe invention and does not pose a limitation on the scope of theinvention otherwise claimed. No language in the specification should beconstrued as indicating any non-claimed element essential to thepractice of the invention. Unless a contrary meaning is explicitlystated, all ranges are inclusive of their endpoints, and open-endedranges are to be interpreted as bounded on the open end by commerciallyfeasible embodiments.

Groupings of alternative elements or embodiments of the inventiondisclosed herein are not to be construed as limitations. Each groupmember can be referred to and claimed individually or in any combinationwith other members of the group or other elements found herein. One ormore members of a group can be included in, or deleted from, a group forreasons of convenience and/or patentability. When any such inclusion ordeletion occurs, the specification is herein deemed to contain the groupas modified thus fulfilling the written description of all Markushgroups used in the appended claims.

It should be apparent to those skilled in the art that many moremodifications besides those already described are possible withoutdeparting from the inventive concepts herein. The inventive subjectmatter, therefore, is not to be restricted except in the spirit of theappended claims. Moreover, in interpreting both the specification andthe claims, all terms should be interpreted in the broadest possiblemanner consistent with the context. In particular, the terms “comprises”and “comprising” should be interpreted as referring to elements,components, or steps in a non-exclusive manner, indicating that thereferenced elements, components, or steps may be present, or utilized,or combined with other elements, components, or steps that are notexpressly referenced. Where the specification claims refers to at leastone of something selected from the group consisting of A, B, C . . . andN, the text should be interpreted as requiring only one element from thegroup, not A plus N, or B plus N, etc.

1. A fleet of aircraft, comprising: a first tiltrotor aircraft having a first rotor with rotor diameter of between 12 and 80 feet inclusive, inboard wing span between 21 and 185 feet inclusive, and fuselage length between 20 and 100 feet inclusive; a second tiltrotor aircraft having a second rotor with rotor diameter of between 12 and 80 feet inclusive, inboard wing span between 21 and 185 feet inclusive, and fuselage length between 20 and 100 feet inclusive; wherein the rotor diameter of the second rotor is at least 10% larger than the rotor diameter of the first rotor; wherein each of the first and second rotors has multiple blades, and a radius measured from a center of rotor rotation, and wherein at 10% of the rotor radius, the flap stiffness of the blades in lbs-in2 is not less than 25 times the diameter of the rotor in feet to the fourth power; and wherein the first and second aircraft having (a) at least 80% commonality by part count of the “green aircraft” or (b) at least 90% commonality in power source and drivetrain.
 2. The fleet of claim 1, wherein a smallest and largest folded length dimension of the first and second aircraft differ by between 15% and 50% of the first aircraft, inclusive.
 3. The fleet of claim 1, wherein a smallest and largest manufacturers empty weight of the first and second aircraft differ by between 15% and 50% of the first aircraft, inclusive.
 4. The fleet of claim 1, wherein a smallest and largest payload length dimension of the first and second aircraft differ by between 15% and 50% of the first aircraft, inclusive.
 5. The fleet of claim 1, wherein the only significant parts difference in power source and drivetrain components between the first and second aircraft is a single gearset.
 6. The fleet of claim 1, wherein at 10% of the rotor radius of each of the first and second rotors, the flap stiffness of each of the blades of each of the first and second rotors in lbs-in2 is not less than 10 times the diameter of the rotor in feet to the fourth power.
 7. The fleet of claim 1, wherein for each of the first and second rotors, a total weight of each blade in lbs. does not exceed 0.0015 times the diameter of the rotor in feet cubed.
 8. The fleet of claim 1, wherein the rotor diameter of the second rotor is at least 20% larger than the rotor diameter of the first rotor.
 9. A method for accommodating differences in rotor diameters in a fleet of tilt-wing rotorcraft having (a) at least 80% commonality by part count of the “green aircraft” or (b) at least 90% commonality by part count in power source and drivetrain commonality, comprising: designing the first aircraft to have a largest rotor diameter that is at least 10% larger than a largest rotor diameter of the second aircraft, and the first aircraft to have a payload weight capacity between 500 and 50,000 pounds, inclusive, greater than a payload capacity of the second aircraft; accommodating the different rotor diameters and payload weight capacities of the first and second aircraft, respectively, by: (a) designing each of the blades of each of the first and second rotors to have flap stiffness of the blades in lbs-in² is not less than 25 times the diameter of the rotor in feet to the fourth power, at 10% of the rotor radius; and (b) designing each of the first and second rotors to be powered through first and second cross wing driveshafts and first and second speed selectable gearboxes, respectively.
 10. The fleet of claim 9, further comprising designing the first and second aircraft wherein the only significant parts difference in power source and drivetrain components between the first and second aircraft is a single gearset.
 11. The method of claim 9, further comprising designing each of the first and second aircraft to sustain in-flight operation over a range of 40% to 100% of maximum rotor RPM over a continuous 20 minute period.
 12. The method of claim 9, further comprising designing each of the first and second aircraft to sustain in-flight operation with at least 25% reduction in RPM between hover and forward speed.
 13. The method of claim 9, wherein each of the blades of each of the first and second rotors has a blade root and a blade tip, and further comprising designing each of the blades of each of the first and second rotors such that the blade stiffness in flap, lag and torsion are continuously decreasing from the blade root to the blade tip. 